This invention relates to gas turbine engine exhaust nozzles and, more particularly, to high performance exhaust nozzles of the variable area variety.
The invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the United States Department of the Air Force.
Exhaust systems are provided for gas turbine engines to direct the rapidly moving exhaust gases rearward into the atmosphere at a velocity and density necessary to produce the required thrust. The advent of new aircraft capable of flying advanced missions, with attendant high aerodynamic loading, requires the development of new propulsion system cycles, and has led to the need for new and unique exhaust nozzle systems. The requirements of the aircraft mission cannot be overlooked in the design of the exhaust system since the mission has a direct and formidable impact on both the performance and mechanical design requirements of the exhaust nozzle system. In particular, the mission impacts the nozzle design in two major areas: exhaust systems performance distribution throughout the flight envelope and exhaust system weight.
The performance of an exhaust nozzle is dictated, to a large extent, by the exhaust nozzle area. The choice of nozzle area is determined by the temperature, mass airflow, velocity and pressure of the exhaust gases. Where the operating range of a gas turbine engine is relatively narrow, the area is optimized at the time of manufacture, and the minor benefits obtainable in performance by providing a variable area capability are offset by increased weight and complexity penalties. However, in modern, high performance engines with broad operating ranges, noise, thrust and fuel economy benefits may be achieved by use of variable area nozzles. In particular, nozzle performance and efficiency is dependent upon matching nozzle throat area (minimum flow area) and nozzle exit area as a function of pressure ratio across the nozzle.
The weight of an exhaust nozzle is dictated, to a large extent, by the complexity of the system in response to an attempt to improve nozzle performance and the structural hardware required to maintain integrity due to loading created by both the exhaust stream and high performance aircraft maneuvering. High maneuver loads by the aircraft cause large pressure differentials and, thus, high nozzle structural loading. To counteract these loadings, the exhaust system's structure and actuator systems require increased strength and, thus, increased weight. Clearly, in aircraft systems where weight is a paramount design consideration, the lightest weight exhaust system is desirable. However, in order for an exhaust system to be a viable concept, it must be mechanically feasible. By this, it is meant that the system must integrate well with both the engine and airframe, must act as a sufficient pressure vessel, and must provide for realistic actuator systems to avoid weight penalties and mechanical instabilities.
It becomes apparent, therefore, that a dichotomy exists when considering exhaust nozzle designs for advanced high performance aircraft since high multimission performance is only attainable with complicated and, thus, relatively heavy exhaust systems. Resolution of this problem requires extensive analysis and iteration of the aircraft mission requirements, the engine performance cycle, and the possible exhaust nozzle configurations.
The requirements of anticipated aircraft missions preclude the utilization of conventional nozzle systems. The convergent nozzle which is commonly used for subsonic flight loses its efficiency because the exhaust stream velocity cannot exceed sonic velocity (Mach number equals one). The convergent-divergent exhaust nozzle permits controlled expansion and acceleration of the exhaust gases after they reach sonic velocity, but these nozzles have a very narrow optimum operating range and must be designed as variable area nozzles to compensate for this characteristic. While such variable area nozzles have been considered in the past, heretofore no exhaust nozzle schemes have been found to be satisfactory for adaptation to a wide range of anticipated future aircraft missions using duct burning turbofan engines.
The problem becomes compounded in engines of the multiple bypass type wherein, in general, the number of nozzles is equivalent to the number of flow ducts within the engine. This is necessitated by the large differences in flow properties between each stream for most engine cycles which generally makes it impractical to combine these flows within the ducts. However, as the number of nozzles increases, so does the engine weight. Therefore, a system is required which provides flow modulation between multiple ducts, which is of relatively simple and lightweight design, and which provides the necessary area variation throughout the mission cycle.